The present invention relates to a missile control system, and more particularly, to a thrust vector control vane and system mounted in the aft portion of a missile and used for steering the missile during launch, as well as a method of making and a method of using such a control vane.
Aircraft and offensive missiles, including some cruise missiles, for example, often fly at low altitudes to avoid detection by enemy radar. In such situations the target, such as a ship, may have only a few seconds to both identify the threat posed by the incoming aircraft or missile and take countermeasures, such as firing a defensive missile. Thus, it is desirable to have a defensive missile locate and disable the offensive missile very quickly.
Land or ship borne defensive missiles generally are launched from a canister in a substantially vertical orientation. Missiles generally have steering control systems which include external aerodynamic control surfaces for guiding the missile. Before its aerodynamic control surfaces or fins are able to affect any significant maneuvers, the missile must achieve a certain minimum velocity, referred to herein as the aerodynamic control velocity, to cause enough air to flow over the aerodynamic control surfaces and provide aerodynamic control. For a ballistic launch trajectory, the missile reaches an altitude of thousands of feet before the aerodynamic control surfaces can cause the missile to pitch over and begin seeking the incoming missile threat. As a result a ballistic launch trajectory is inefficient, time consuming, and limits the missile sensor line-of-sight capabilities for optimum target detection and tracking.
A number of systems have been developed in an attempt to maneuver the missile prior to reaching an aerodynamic control velocity and to decrease reaction time after sensing an incoming threat. However, most current devices, although generally acceptable for some uses, have been found to be inadequate for many applications. Detachable jet tab systems, for example, formed of auxiliary propulsion units mounted to missile fins conflict with folding control surfaces. Folding control surfaces generally are necessary for any missile to be loaded into a launch canister having stringent volume constraints. Detachable jet tab systems require increases in the launch canister cross-sectional area for additional volume taken up by the jet tabs external to the missile fuselage.
Existing systems generally also can be classified as either nondetachable or ejectable, the latter often incorporating redundant control electronics. Nondetachable systems limit mission range and performance with rocket thrust degradation throughout the missile trajectory. Detachable, self-actuation mechanisms are substantially heavier and inherently more complicated than nondetachable systems. The increased complexity leads to reduced reliability, and the added weight requires more rocket propellant for missile launch and flight to the target. An actively detachable system generally uses a pyrotechnic actuated ejection mechanism and disengageable power coupling drive that introduces weight, complexity, and multiple operational failure risks. Furthermore, the act of ejecting the control system can knock the missile off its intended trajectory.
To overcome the deficiencies of prior systems, systems have been developed that place a mechanism in the exhaust plume of the rocket engine for control purposes, providing control immediately upon launch. Generally, the purpose is to pitch the missile over (rotate the missile about an axis transverse to the longitudinal axis and previous direction of flight during launch) and to avoid rolling. Rolling generally interferes with operation of the missile guidance system and is a problem that is minimized at low velocities by placing the control surfaces within the exhaust plume.
So-called xe2x80x9cerodiblexe2x80x9d control surfaces have been developed that are placed in the path of rocket engine exhaust and break apart after a period of time. However, these often break apart in larger pieces than generally is acceptable. It would be desirable to avoid ejecting large pieces of material from the missile during launch and flight.
The present invention provides a thrust vector control jet vane, a jet vane control system and a missile incorporating such jet vanes, as well as a method of making such jet vanes. The jet vanes form aerodynamic surfaces for generating vehicle maneuvering forces by diverting the propellant plume at missile launch. As an aerodynamic control velocity is achieved and the aerodynamic control surfaces external to the vehicle airframe assume command authority, the jet vanes dissolve into granular particulates in the propellant plume. Therefore, rocket motor propulsion efficiency or specific impulse is not degraded beyond that required to perform launch maneuvers, reaction time is decreased, and the missile exhibits improved kinetic performance during the aerodynamic control phase of the powered flight to the target.
Dissolvable jet control vanes provide numerous advantages in the design, construction and/or performance of a missile. For example, dissolvable control vanes eliminate or minimize the often tortuous practice of determining the least desirable inefficiency to the overall missile system. The control vanes gracefully disintegrate in a timely basis, providing a disposable control vane for maneuvering the missile for pitch, yaw and roll control immediately upon launch. Rocket motor propulsion efficiency or specific impulse is not degraded beyond that required to perform the launch maneuvers, hence the missile exhibits improved kinetic performance during the powered flight phase to the target. Interception of highly mobile targets at an extended range is further enhanced by dissolvably jettisoning the control vanes after pitch-over. Ease of thrust vane control operation without the activation of pyrotechnic-actuated ejection mechanisms and greater reliability resulting from system simplification are additional advantages that also lead to cost and risk reduction.
Dissolvable control vanes are possible through the utilization of multiple advanced composite materials designed to perform different individual functions on a time limited basis, yet integrated or colaminated together to achieve a combined, pre-programmed structural capability by taking advantage of their known high temperature performance characteristics and environmental limitations. As a result, failure of the composite control vanes produced in accordance with the present invention can be precisely controlled in a manner unforeseen in prior xe2x80x9cerodiblexe2x80x9d material designs. The composite dissolvable jet vanes also provide an inexpensive, disposable thrust vector control methodology for retrofitting high performance missiles for low speed surface launch applications with thrust vector control requirements.
The dissolvable jet vane provides a lightweight, reliable means of removing steering jet vanes from the exhaust stream of a solid rocket motor nozzle. The dissolvable jet vane materials withstand the pressure and thermal loads associated with missile steering during the first few seconds of rocket boost until the missile obtains sufficient speed to use conventional external aerodynamic control surfaces for steering control. Once control passes to the external fin, the jet vanes rapidly and uniformly dissolve in the exhaust stream.
According to one aspect of the invention, a dissolvable thrust vector control vane includes a frame and a thermal protection layer on at least a portion of the frame. In accordance with one embodiment of the invention, the dissolvable control vane further includes an erosion-resistant material on at least a forward edge of the frame. The erosion-resistant material forms an insert that is mounted to the forward edge of the frame; the insert includes a carbon-carbon structure and a surface coating on the structure selected from a group including a ceramic, a carbide, and a metallic material; the surface coating includes one or more materials selected from the group including a hafnium-diboride ceramic, a zirconium-diboride ceramic, a hafnium-carbide, a tantalum-carbide, or a metallic rhenium refractory coating; the surface coating generally has a thickness of less than about five thousandths of an inch (0.127 mm) thick; the insert further includes one or more materials selected from a group including ceramic and oxide-based oxygen permeability and volatility barriers; the insert has a generally T-shape cross-section with the tail of the T connected to the frame; and/or the insert has a generally U-shape cross-section with the open end of the U attached to the frame.
According to one or more embodiments of the invention, the frame includes graphite-reinforced organic resins; the graphite-reinforced organic resins are carbon-reinforced, high temperature organic resins that include one or more resins selected from a group including: PMR-15 based polyimides, PT cyanate esters, bismaleimides, phthalonitriles, and Avimid N; the thermal protection layer includes an ablative insulator; the thermal protection layer includes one or more materials selected from a group including chopped quartz, carbon, and silica fiber-reinforced phenolic resins; the control assembly includes a control shaft connected to the frame; the control shaft is formed from a refractory material; and/or the frame is generally planar, and the control vane further comprises an integral blast shield extending in a plane that is generally perpendicular to the plane of the frame.
The dissolvable control vane of another embodiment further includes a control assembly connected to the frame for controlling the orientation of the vane.
In accordance with another aspect of the invention, a thrust vector control system for a missile having a motor for propelling the missile that creates an exhaust plume, including a dissolvable control vane and a control assembly for controlling the orientation of the control vane. The system is mountable to the missile such that the control vane is within a path of the exhaust plume.
In accordance with another aspect of the invention, a missile includes a motor for propelling the missile that creates an exhaust plume, and a dissolvable control vane mounted within a path of the exhaust plume.
In accordance with yet another aspect of the invention, a method of steering a missile during launch includes the steps of placing a movable control vane in a path of an exhaust plume of the missile, launching the missile, and controllably moving the vane to steer the missile. The step of launching the missile includes igniting the motor to expel the exhaust plume and dissolving the vane within the exhaust plume within a predetermined period of time.
In accordance with one or more embodiments of the invention the step of moving the control vane includes moving the control vane to provide pitch, yaw and roll control.
In accordance with yet another aspect of the invention, a method of making a dissolvable missile control vane includes forming a frame from graphite-reinforced organic resins and applying a thermal protection layer to the frame.
According to one or more embodiments of the invention, the step of forming the frame includes using high temperature organic resins selected from a group including PMR-15 based polyimides, PT cyanate esters, bismaleimides, phthalonitriles, and Avimid N; the step of applying the thermal protection layer includes laminating an ablative insulator onto the frame; the step of applying the thermal protection layer includes laminating one or more materials selected from a group of chopped quartz, carbon and silica fiber-reinforced phenolic resins; the step of applying the thermal protection layer includes compression molding; applying the thermal protection layer includes wrapping continuous fiber-reinforced laminates onto the frame; the step of applying the thermal protection layer includes mechanically interlocking the thermal protection layer to the frame by drilling a plurality of holes into the frame and molding the thermal protection layer into the holes for additional adhesive attachment; the step of forming the frame includes attaching a metal shaft to the frame; and/or the step of forming the thermal protection layer includes forming the thermal protection layer over at least a portion of the shaft.
One or more embodiments may further include the step of mounting an erosion-resistant material at a forward edge of the frame; wherein the step of mounting an erosion-resistant material includes forming a carbon-carbon insert; the step of forming the carbon-carbon insert includes applying a coating formed from a material selected from a group including zirconium-diboride ceramics and hafnium-diboride ceramics, reinforced with a material selected from a group including silicon-carbide and ultra-high temperature ceramics; the step of applying the coating includes using vapor deposition; the step of applying the coating includes using a fused slurry process; the step of applying the coating includes forming a coating which generally is less than about five thousandths of an inch (0.127 mm) thick; the step of applying the coating includes using an adherent reaction-sintered material as a transitional bond layer; in the step of using the transitional bond layer includes using a material selected from a group including hafnium-carbide and hafnium-diboride; in the step of mounting an erosion-resistant material further includes applying a coating selected from a group including an oxygen permeability barrier and a volatility barrier; in the step of applying the coating includes applying a coating selected from a group including ceramic and oxide-based coatings; in the step of forming the insert includes using a material selected from a group including zirconium-diboride ceramics and hafnium-diboride ceramics that is reinforced with silicon-carbide fibers; the step of forming the insert includes using an ultra-high temperature ceramic; the step of forming includes forming an insert having a generally T-shape cross-section with the tail of the T connected to the frame; and/or the step of forming includes forming an insert having a generally U-shape cross-section with the open end of the U attached to the frame.
According to still another aspect of the invention, a thrust vector control vane, includes the combination of a frame for temporarily providing structural support, a thermal protection laminate applied at least to the frame for temporarily thermally isolating the frame, and an erosion-resistant insert connected to a leading edge of the frame for temporarily shielding the frame from particulate impacts.
The foregoing and other features of the invention are hereinafter fully described and particularly pointed out in the claims, the following description and annexed drawings setting forth in detail a certain illustrative embodiment of the invention, this embodiment being indicative, however, of but one of the various ways in which the principles of the invention may be employed.